1. Technical Field
The present invention relates in general to inlet design for aircraft engines and, in particular, to an improved system, method, and apparatus for throat corner scoop offtake for mixed compression inlets for high speed aircraft engine applications.
2. Description of the Related Art
Air inlet systems for gas turbine powered supersonic aircraft are required to decelerate the approaching flow to subsonic conditions before it reaches the engine face. Supersonically, this can be done through shock waves or isentropic compression generated externally, internally, or by a mixture of both. Fixed geometry external compression inlets have typically been used for aircraft (e.g., the F-16 and F-18) designed for short excursions to supersonic conditions, due to the relative simplicity and light weight of these designs. Aircraft capable of higher speeds, such as the F-14 and F-15, have employed variable geometry external compression inlets to obtain better engine and inlet airflow matching at low speeds, and higher performance at supersonic speeds.
High altitude supersonic cruise aircraft typically require maximum efficiency at the cruise point to obtain optimum range and payload. At speeds above Mach 2, mixed compression inlet systems become favorable over external compression systems due to reduced drag. Mixed compression inlets have been demonstrated in flight on aircraft such as the A-12, SR-71, D-21, and XB-70. Several other designs have been tested over the past 50 years. All of these mixed compression designs were based on either axi-symmetric or two-dimensional compression schemes in order to minimize shock interactions caused by complex, three-dimensional geometry.
As shown in FIG. 1, axi-symmetric mixed compression inlet designs 11 typically include a throat bleed system that removes the low pressure boundary layer from the main duct 13. This provides terminal normal shock stability and reduces shock/boundary layer interaction, which reduces overall pressure recovery and increases distortion. In this example, the throat bleed system includes both a centerbody shock trap 15 and a cowl slot 17. The low energy air captured in the shock trap 15 would likely be exhausted overboard as it typically does not have enough energy to be used as utility flow. Having more energy due to a larger dynamic pressure component, the cowl slot 17 could possibly be used for utility flow. Various approaches to these bleed systems have been implemented in the industry for axi-symmetric and two-dimensional mixed compression inlets. Increasing demand for more integrated inlet and airframe concepts has resulted in the need for more exotic inlet aperture shapes. These exotic shapes impose additional geometric constraints that require novel approaches to bleed system design and integration.
Exotically-shaped, high speed engine inlets can suffer from several diverse performance losses. First, mixed compression inlets with duct wall interfaces that form acute angles (such as streamline traced inlets) can develop vorticity and a thick boundary layer (e.g., corner flow) in these regions which can cause separation and flowfield distortion that reduces engine performance. Second, a mixed compression inlet can undergo a process called “unstart” in which terminal shock stability is lost and airflow to the engine is drastically reduced, which consequently reduces engine performance. Third, airflow from the engine inlet is required for nacelle ventilation, environmental control systems, and various other utility and subsystems. While current state-of-the-art bleed system designs and integration approaches are workable for axi-symmetric and two-dimensional mixed compression inlets, an improved solution would be desirable for advanced shaped mixed compression inlet concepts that impose additional requirements based on geometrical constraints.